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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 127 records · Page 7
All sourcesNASA NTRSDOE OSTI
NASA NTRS1974

Effect of damper on overall and blade-element performance of a compressor rotor having a tip speed of 1151 feet per second and an aspect ratio of 3.6

The overall and blade-element performance of two configurations of a moderately high aspect ratio transonic compressor rotor are presented. The subject rotor has conventional blade dampers. The performance is compared with a rotor utilizing dual wire friction dampers. At design speed the subject achieved a pressure ratio of 1.52 and efficiency of 0.89 at a near design weight flow of 72.1 pounds per second. The rotor with wire dampers gave consistently higher pressure ratios at each speed, but efficiencies for the two rotors were about the same. Stall margin for the subject rotor was 20.4 percent, but for the wire damped rotor only 4.0 percent.

Lewis, G. W.↗
NASA NTRS1974

Design and performance of a high-pressure-ratio, highly loaded axial-flow transonic compressor space

A 50-cm-diam. axial-flow transonic compressor stage with multiple-circular-arc blades was designed and tested. At design speed, a rotor peak efficiency of 0.85 occurred at an equivalent weight flow of 29.3 kg/sec. Stage peak efficiency was 0.79 at 28.6 kg/sec. Stage total pressure ratio at peak efficiency was 1.84. The stall margin at design speed was 5 percent. Rotor and stator losses were higher than predicted. The stator choked at design flow.

Lewis, G. W., Jr.↗
NASA NTRS1974

Sealing technology for aircraft gas turbine engines

Experimental evaluation under simulated engine conditions revealed that conventional mainshaft seals have disadvantages of high gas leakage rates and wear. An advanced seal concept, the self-acting face seal, has a much lower gas leakage rate and greater pressure and speed capability. In endurance tests (150 hr) to 43 200 rpm the self-acting seal wear was not measurable, indicating noncontact sealing operation was maintained even at this high rotative speed. A review of published data revealed that the leakage through gas path seals has a significant effect on TSFC, stall margin and engine maintenance. Reducing leakages by reducing seal clearances results in rubbing contact, and then the seal thermal response and wear determines the final seal clearances. The control of clearances requires a material with the proper combination of rub tolerance (abradability) and erosion resistance. Increased rub tolerance is usually gained at the expense of reduced erosion resistance and vice versa.

Ludwig, L. P.↗
NASA NTRS1974

Transonic compressor technology advancements

The highlights of the NASA program on transonic compressors are presented. Effects of blade shape and throat area on losses and flow range are discussed. Some effects of casing treatment on stall margin are presented. Results of tests with varying solidity are also presented. High Mach number, highly loaded stators are discussed and some results of stator hub slit suction are presented.

Benser, W. A.↗
NASA NTRS1974

Aerodynamic performance of 0.5 meter-diameter, 337 meter-per-second tip speed, 1.5 pressure-ratio, single-stage fan designed for low noise aircraft engines

Overall and blade-element aerodynamic performance of a 0.271-scale model of QF-1 are presented, examined, and then compared and evaluated with that from similar low noise fan stage designs. The tests cover a wide range of speeds and weight flows along with variations in stator setting angle and stator axial spacing from the rotor. At design speed with stator at design setting angle and a fixed distance between stage measuring stations, there were no significant effects of increasing the axial spacing between rotor stator from 1.0 to 3.5 rotor chords on stage overall pressure ratio, efficiency or stall margin.

Gelder, T. F.↗
NASA NTRS1975

Digital implementation of the TF30-P-3 turbofan engine control

The standard hydromechanical control modes for TF30-P-3 engine were implemented on a digital process control computer. Programming methods are described, and a method is presented to solve stability problems associated with fast response dynamic loops contained within the exhaust nozzle control. A modification of the exhaust nozzle control to provide for either velocity or position servoactuation systems is discussed. Transient response of the digital control was evaluated by tests on a real time hybrid simulation of the TF30-P-3 engine. It is shown that the deadtime produced by the calculation time delay between sampling and final output is more significant to transient response than the effects associated with sampling rate alone. For the main fuel control, extended update and calculation times resulted in a lengthened transient response to throttle bursts from idle to intermediate with an increase in high pressure compressor stall margin. Extremely long update intervals of 250 msec could be achieved without instability. Update extension for the exhaust nozzle control resulted in a delayed response of the afterburner light-off detector and exhaust nozzle overshoot with resulting fan oversuppression. Long update times of 150 msec caused failure of the control due to a false indication by the blowout detector.

Cwynar, D. S.↗
NASA NTRS1974

Sealing technology for aircraft gas turbine engines

Experimental evaluation under simulated engine conditions revealed that conventional mainshaft seals have disadvantages of high gas leakage rates and wear. An advanced seal concept, the self-acting face seal, has a much lower gas leakage rate and greater pressure and speed capability. In endurance tests (150 hr) to 43,200 rpm the self-acting seal wear was not measurable, indicating that noncontact sealing operation was maintained even at this high rotative speed. A review of published data revealed that the leakage through gas path seals has a significant effect on thrust specific fuel consumption, stall margin, and engine maintenance. Reducing leakages by reducing seal clearances results in rubbing contact, and then the seal thermal response and wear determines the final seal clearances.

Ludwig, L. P.↗
NASA NTRS1976

Performance of low-pressure-ratio low-tip-speed fan stage with blade tip solidity of 0.65

The overall and blade-element performance of a low pressure ratio, low tip speed fan stage is presented over the stable operating range at rotative speeds from 90 to 120 percent of design speed. Stage peak efficiency of 0.927 was obtained at a weight flow of 32.4 kg/sec (190.31 kg/sec/sq m of annulus area) and a pressure ratio of 1.134. The stall margin at design speed and peak efficiency was 15.3 percent.

Kovich, G.↗
NASA NTRS1976

Effect of casing treatment on performance of an inlet stage for a transonic multistage compressor

An inlet stage of a transonic compressor was tested with three rotor tip casing treatment configurations: blade angle slots, circumferential grooves, and axial skewed slots. Significant increases in both rotor and stage total pressure ratio, total temperature ratio, efficiency, flow range, and very large improvements in stall margin were obtained with all three casing treatment configurations. The greatest improvement in performance was achieved with axial skewed slots.

Urasek, D. C.↗
NASA NTRS1976

Altitude performance of a low-noise-technology fan in a turbofan engine with and without a sound suppressing nacelle

Test variables were inlet Reynolds number index (0.2 to 0.5), flight Mach number (0.2 to 0.8), and flow distortion (tip radial and combined circumferential - tip radial patterns). Results are limited to fan bypass and overall engine performance. There were no discernible effects of Reynolds number on fan performance. Increasing flight Mach number shifted the fan operating line such that pressure ratio decreased and airflow increased. Inlet flow distortion lowered stall margin. For a Reynolds number index of 0.2 and flight Mach number of 0.54, the sound suppressing nacelle lowered fan efficiency three points and increased specific fuel consumption about 10 percent.

Biesiadny, T. J.↗
NASA NTRS1976

Single stage, low noise, advanced technology fan. Volume 5: Fan acoustics. Section 1: Results and analysis

The acoustic tests and data analysis for a 0.508-scale fan vehicle of a 111,300 newton (25,000 pound) thrust, full-size engine, which would have application on an advanced transport aircraft, is described. The single-stage advanced technology fan was designed to a pressure ratio of 1.8 at a tip speed of 503 m/sec (1,650 ft/sec) to achieve the desired pressure ratio in a single-stage fan with low radius ratio (0.38), and to maintain adequate stall margin. The fan has 44 tip-shrouded rotor blades and 90 outlet guide vanes. The two basic approaches taken in the acoustic design were: (1) minimization of noise at the source, and (2) suppression of the generated noise in the inlet and bypass exhaust duct. Suppression of the generated noise was accomplished in the inlet through use of the hybrid concept (wall acoustic treatment plus airflow acceleration suppression) and in the exhaust duct with extensive acoustic treatment including a splitter. The goal of the design was attainment of twenty effective perceived noise decibels (20 EPNdB) below current Federal Air Regulation noise standards for a full-scale fan at the takeoff, cutback, and approach conditions. The suppression goal of FAR 36-20 was not reached, but improvements in the technology of both front and aft fan-noise suppression were realized. The suppressed fan noise was shown to be consistent with the proposed federal regulation on aircraft noise.

Jutras, R. R.↗
NASA NTRS1976

Single stage, low noise, advanced technology fan. Volume 5: Fan acoustics. Section 2: One-third octave data tabulations and selected narrowband traces

The raw-acoustic data corrected to standard day, from acoustic tests performed on a 0.508-scale fan vehicle of a 111,300 newton thrust, full-size engine, which has application on an advanced transport aircraft, are presented. The single-stage advanced technology fan was designed to a pressure ratio of 1.8 at a tip speed of 503 m/sec to achieve the desired pressure ratio in a single-stage fan with low radius ratio, and to maintain adequate stall margin. The two basic approaches taken in the acoustic design were: (1) minimization of noise at the source, and (2) suppression of the generated noise in the inlet and bypass exhaust duct. Suppression of the generated noise was accomplished in the inlet through use of the hybrid concept (wall acoustic treatment plus airflow acceleration suppression) and in the exhaust duct with extensive acoustic treatment including a splitter. The goal of the design was attainment of twenty effective perceived noise decibels. The suppression goal of FAR 36-20 was not reached, but improvements in the technology of both front and aft fan-noise suppression were realized.

Jutras, R. R.↗
NASA NTRS1978

Gas Path Sealing in Turbine Engines

A survey of gas path seals is presented with particular attention given to sealing clearance effects on engine component efficiency. The effects on compressor pressure ratio and stall margin are pointed out. Various case-rotor relative displacements, which affect gas path seal clearances, are identified. Forces produced by nonuniform sealing clearances and their effect on rotor stability are discussed qualitatively, and recent work on turbine-blade-tip sealing for high temperature is described. The need for active clearance control and for engine structural analysis is discussed. The functions of the internal-flow system and its seals are reviewed.

Ludwig, L. P.↗
NASA NTRS1978

Performance with and without inlet radial distortion of a transonic fan stage designed for reduced loading in the tip region

A transonic compressor stage designed for a reduced loading in the tip region of the rotor blades was tested with and without inlet radial distortion. The rotor was 50 cm in diameter and designed for an operating tip speed of 420 m/sec. Although the rotor blade loading in the tip region was reduced to provide additional operating range, analysis of the data indicates that the flow around the damper appears to be critical and limited the stable operating range of this stage. For all levels of tip and hub radial distortion, there was a large reduction in the rotor stall margin.

Schmidt, J. F.↗
NASA NTRS1977

Optimal controls for an advanced turbofan engine

Linear optimal control theory is applied to the control synthesis of a high bypass ratio, variable pitch, turbofan engine. The basic control philosophy is to use only a low order dynamic model of the plant coupled with the concept of integral-output states so as to maintain control simplicity yet guarantee integral control of thrust, turbine temperature and other important engine outputs. Linear simulation results indicate that the control system developed provides rapid control of small thrust perturbations and quickly eliminates the effect of unmodelled thrust and temperature disturbances. Large thrust accelerations are obtained in about one half second while the control maintains negligible overshoot in temperature and stall margins.

Slater, G. L.↗
NASA NTRS1978

Gas path sealing in turbine engines

Survey of gas path seals is presented with particular attention given to sealing clearance effects on engine component efficiency. The effects on compressor pressure ratio and stall margin are pointed out. Various case-rotor relative displacements, which affect gas path seal clearances, are identified. Forces produced by nonuniform sealing clearances and their effect on rotor stability are discussed qualitatively, and recent work on turbine-blade-tip sealing for high temperatures is described. The need for active clearance control and for engine structural analysis is discussed. The functions of the internal-flow system and its seals are reviewed.

Ludwig, L. P.↗
NASA NTRS1978

Aerodynamic performance of a 1.35-pressure-ratio axial-flow fan stage

The overall blade element performances and the aerodynamic design parameters are presented for a 1.35-pressure-ratio fan stage. The fan stage was designed for a weight flow of 32.7 kilograms per second and a tip speed of 302.8 meters per second. At design speed the stage peak efficiency of 0.879 occurred at a pressure ratio of 1.329 and design flow. Stage stall margin was approximately 14 percent. At design flow rotor efficiency was 0.94 and the pressure ratio was 1.360.

Osborn, W. M.↗
NASA NTRS1978

Performance of single-stage axial-flow transonic compressor with rotor and stator aspect ratios of 1.19 and 1.26, respectively, and with design pressure ratio of 1.82

The overall and blade-element performance of a low-aspect-ratio transonic compressor stage is presented over the stable operating flow range at 70, 90, and 100 percent design speeds. At design speed the rotor and stage achieved peak efficiencies of 0.872 and 0.845 at pressure ratios of 1.875 and 1.842, respectively. The stage stall margin at design speed was 21.8 percent.

Reid, L.↗